15 results about "Axial compressor" patented technology

This invention discloses a rotor blade of a segmented stacked axial compressor and its design method. The design method includes: determining the axial position of the stacking line; determining the circumferential position of the stacking line: the stacking line is S-shaped in the circumferential direction, and the stacking line includes four control points A, B, C, and D sequentially from the blade root to the blade tip along the height direction; the coordinates of the centroid positions of each basic airfoil section in the first interval from point B to point C conform to a specific formula; and determining the shape of the stacking line based on the determined axial and circumferential positions. This invention, by setting the stacking line of the rotor blade to be S-shaped in the circumferential direction, generates a reverse pressure gradient in the radial direction, causing the gas at both ends to migrate towards the blade center, thereby improving the flow in the end region. This reduces the influence of secondary flow in the blade tip region; and by setting a specific function, the circumferential offset of the stacking line can be controlled by constructing a simple equation, improving the computational efficiency of optimizing the airfoil.

A multi-stage axial compressor comprises a rotating shaft on which a plurality of impeller blades are mounted, a casing surrounding the rotating shaft, the casing forming a flow path of a working fluid between the rotating shaft and the casing, a wall section having an annular shape and extending in a circumferential direction of the rotating shaft to surround the casing, the wall section forming a dispensing chamber having an annular shape and connected to the flow passage, a plurality of connection sections connected to the outer circumferential surface of the wall section, the connection sections forming respective outlet flow passages connected to the dispensing chamber, and a plurality of dispensing pipes connected to the respective connection sections.In a cross-sectional view orthogonal to the rotating shaft, of two corner regions where an inner surface of each of the connection sections and an inner surface of the wall section intersect or cross, at the corner region located on a rear side in a direction of rotation of the working fluid in the dispensing chamber, where θ1 is defined as an angle formed between an inner surface of the connection section and the inner surface of the wall section, the angle 01 is not greater than 225°.

This invention relates to the field of aerodynamic design technology for multi-stage axial compressors, and discloses a design method and system for passive stability enhancement of a multi-stage axial compressor through self-circulation. By calculating the aerodynamic performance of the multi-stage axial compressor and evaluating engine stability, the method identifies operating conditions where the multi-stage axial compressor has insufficient stability margin. Unsteady flow field simulations are then conducted under these conditions to determine the first stage stall. For the first stage rotor tip stall, a self-circulating casing is used for passive stability enhancement of the multi-stage compressor. The self-circulating casing structure is derived from sensitivity analysis of its design parameters, selecting the optimal design parameters to ensure an increase in the stability margin of the multi-stage axial compressor across the entire speed range. This satisfies the stability margin requirements of multi-stage axial compressors and avoids the problem of insufficient stability margin in high-load multi-stage axial compressors across the entire operating speed range.

This invention belongs to the technical field of compressor aerodynamic stability analysis. It proposes a deterministic learning-based method and system for analyzing the aerodynamic stability of axial compressors. Using pulsating pressure data during the aerodynamic instability development process of an axial compressor, a deterministic learning identification algorithm based on radial basis function neural networks is used to model the nonlinear dynamics of the compressor's internal flow field, thereby obtaining dynamic diagrams of normal and instability precursor states. Dynamic indices are extracted from these diagrams to construct dynamic feature vectors describing the compressor's aerodynamic stability. Based on the dynamic feature vectors corresponding to different operating states, the dynamic classification boundary between normal and instability precursor states is determined. The distance between the dynamic feature vector of the compressor data under test and the dynamic classification boundary is calculated to determine the compressor's operating state. This invention enables early warning of aerodynamic instability and has good interpretability and engineering application value.

Provided is an axial compressor comprising an exit guide blade that is provided on the downstream side of the final-stage stationary blade. When, in a cross section orthogonal to the blade height direction of the final-stage stationary blade, an extended straight line of a camber line of the final-stage stationary blade from the trailing edge of the final-stage stationary blade is referred to as a camber line extension line, the shortest distance between the camber line extension line and the negative pressure surface of the exit guide blade is Dmim, and the chord length of the exit guide blade is C, the final-stage stationary blade and the exit guide blade satisfy the expression 0.07 ≤ Dmin / C ≤ 0.20 in the entire range of the final-stage stationary blade in the blade height direction.

Axial compressor (10) with casing treatment for an aircraft turbomachine, comprising an annular casing (12) having at least one annular row of air outlet ports (20) and at least one annular row of air inlet ports (22), the air inlet ports (22) being connected to the air outlet ports (20) by channels (24) allowing air recirculation from the air inlet ports (22) to the air outlet ports (20), the compressor (10) further comprising a shell (32) which is housed in a slot (26) of the casing (12) and which has openings (34), this shell (32) being radially extendable and rotationally movable in the slot (26). Figure for the abbreviation: Figure 1

The present invention is directed to axial-centrifugal contra-rotating compressor system comprising combination of contra-rotating axial compressor stage with a back end centrifugal compressor stage, eliminating stator blades, an inter-stage connecting duct at the entry of the centrifugal compressor and an discharge diffuser at the exit of centrifugal compressor rotor characterized by radial to axial turning annular passage as per the gas turbine engine requirement, favouring reduction in engine size, weight, number of parts, overall engine length and produces a higher-pressure rise and swirl-free discharge towards the combustion chamber. The first low-pressure rotor having at least one axial blading assembly rotates in the direction opposite to the second high-pressure rotor blading assembly comprising of at least one axial rotor and one centrifugal rotor. The blades for axial compressor are adopted of low-aspect ratio configuration for reliable high load stall free operation. The inter-stage connecting duct minimizes total pressure loss. The discharge diffuser recovers higher exit velocity and hence higher kinetic energy.

An adjustable bypass ratio low-bypass turbofan engine comprises a fan, an outer bypass duct, an axial compressor, a combustion chamber, a turbine, and an exhaust nozzle. The fan has three stages. The first two stages have blades of equal length, with an airflow ratio to the inner bypass duct not exceeding 1.3. The third stage blades are longer than the first two stages. Adjustable guide vanes are also present, either behind the second-stage blades or in front of or behind the portion of the third-stage blades that are longer than the first two stages. During startup, by opening or closing the adjustable guide vanes, the flow area of ​​the third-stage fan blades is changed, or the airflow of the outer bypass duct is adjusted. This allows the engine to operate at a maximum bypass ratio of 0.3 or higher but not exceeding 1.0 during subsonic cruise, and at a minimum bypass ratio of 0.3 or less during supersonic flight. This improves the engine's thrust and fuel efficiency at subsonic and supersonic speeds, but without requiring a core drive fan system or a three-bypass duct.

This application provides a support base for the stator vane adjusting ring of an ultra-large axial compressor. The support base is disposed between the housing and the adjusting ring to support the adjusting ring. The support base includes: a support frame disposed on the housing, with rollers disposed on the support frame and the rollers contacting the adjusting ring; and a fixing member passing through the support frame and connected to the housing to assemble the support frame onto the housing. This application, through the inclusion of a buffer assembly, allows the buffer assembly to undergo adaptive deformation under pressure when the housing expands. This deformation, combined with a buffer gap, absorbs and offsets the expansion, preventing excessive supporting stress due to thermal expansion differences. This effectively protects the thin-walled housing and adjusting ring of the ultra-large axial compressor from damage, while maintaining the concentricity of the adjusting ring and the housing.

A method for predicting leakage losses in axial compressor grates is disclosed. This method calculates the grate leakage flow rate based on the inlet and outlet aerodynamic parameters at the compressor's planar blade mid-diameter and the grate geometric parameters. Then, it calculates the compressor's mainstream loss caused by the grate leakage by comparing the grate leakage flow rate with the compressor's mainstream flow rate. Finally, it determines the influence range and radial distribution law of the grate leakage loss using the compressor's planar blade geometric parameters. This invention targets compressor stators with grate seals. During compressor design and analysis, it eliminates the need for experimental measurements. It predicts grate leakage flow losses using the aerodynamic parameters at the stator's mid-diameter and the geometric parameters of the grate cavity, considering the impact of the leakage flow on the compressor stator blade passages. This allows for the incorporation of the radial influence of grate leakage on compressor outlet performance into the compressor's flow path design and analysis, predicting the grate leakage flow rate, and the influence range and magnitude of the stator outlet loss caused by the grate leakage flow, thus achieving more refined flow path calculations for axial compressors.

A propulsion system includes a fan rotor connected to a fan shaft, a first spool including a first turbine that drives the fan rotor via a first shaft, and a second spool including a second turbine and a second axial compressor. The second turbine is two-stage and drives the second compressor via a second shaft. The second shaft rotates at a higher speed than a rotation speed of the first shaft, and a hub-to-tip ratio at an inlet of the second compressor is between 0.41 and 0.60. A reduction structure couples the first shaft and the fan shaft to drive the fan shaft at a rotation speed lower than the rotation speed of the first shaft. The second spool has a relationship among an average outer radius of the second compressor, an average outer radius of the second turbine, a number of stages in the second compressor, a several constants.

An integral gear compressor includes a first pinion shaft, a second pinion shaft, and a large gear supported for rotation within a gearbox. A first compressor unit is suspended at a first end of the first pinion shaft, and a second compressor unit is suspended at either the second end of the first pinion shaft, or at either the first or second end of the second pinion shaft. A third compressor unit is suspended at either the first or second end of the second pinion shaft. Both the second and third compressor units are centrifugal compressor units. The first compressor unit includes an axial compressor section and a centrifugal compressor section combined with each other.

The invention relates to a method for avoiding drop impingement erosion in an axial compressor (1), wherein the axial compressor (1) has a rotatably mounted rotor (3) and a housing (5) arranged around the rotor (3), wherein a flow channel (7) is arranged between the rotor (3) and the housing (5), wherein the rotor (3) comprises compressor rotor blades (4), wherein the housing (5) comprises compressor guide blades (6), wherein the flow channel (7) is formed by the compressor guide blades (6) and compressor rotor blades (4), wherein at least one compressor guide blade (6) having a temperature device is designed in such a way that the compressor guide blade (6) can be heated, and the temperature device is designed in such a way that accumulations of a process medium on the surface (13) of the heated compressor guide blade (6) evaporate.

The application relates to the technical field of industrial automation, and particularly provides an axial flow compressor system supporting distributed control, which comprises a controller and a plurality of remote IO modules, one remote IO module is arranged in different unit composition modules of a compressor unit; a first remote IO module corresponding to a shafting equipment sub-module is used for monitoring host equipment body measuring points; a second remote IO module corresponding to an auxiliary machine sub-module is used for monitoring fan inlet and exhaust pipeline measuring points, lubricating oil power oil system measuring points and water system measuring points; a third remote IO module corresponding to a process measuring point sub-module is used for monitoring process pipeline measuring points; and a fourth remote IO module corresponding to an electrical sub-module is used for monitoring electrical system measuring points. The application monitors corresponding unit composition modules through the plurality of remote IO modules, so that the system hardware reliability is improved, the space requirement between centralized control cabinets is reduced, and the arrangement is flexible and strong in adaptability.